text:
What exactly causes pitch moment in the Phugoid mode?
I am trying to figure out exactly what causes the pitching moment for the Phugoid mode? We know that alpha remains mostly constant, so what is causing the pitching moment? I thought maybe it is the trim point of the aircraft, but if trim was exactly neutral then how would the phugoid mode still proceed, since alpha is zero and at the design condition there would be zero tail load?
I guess my question then becomes what derivatives are responsible for causing the Phugoid mode or am I missing somehow that is just the dynamic response to the derivatives as a whole?

"since alpha is zero" -- why?

Everything. When we are talking about flight dynamics and stability in particular, it is the balance of forces or moments that we must consider. For simplicity of analysis, we often choose reference points that allow us to isolate some effects and make them independent of some variables, but if you are interested in the "physics" of it, then you should keep in mind that, in general, the forces and moments are everywhere, and they change.
Specifically, you seem to think that the tail load is zero in the trimmed condition. Not true. It could be true only for aircraft with neutral static longitudinal stability, which normally wouldn't display phugoid oscillations.
In the trimmed conditions, both the tail and the wing create moments (with respect to CG), but they balance out to zero. The essence of static stability is that upon the disturbance, the balance is offset such that it leads to restoration of the initial conditions.
In pure phugoid motion, the upset is airspeed but not angle of attack (AoA). Now, an airplane may have a tendency to have a pitch moment directly in response to speed, which we can designate $M^V$, or better as a coefficient $c_m^V$. This can be related to the nature of its thrust and/or moments from other parts. But this is not strictly required, and we can consider the simplest classical system of just a wing and a fixed tailplane. This system will not see a balance change simply due to airspeed. But consider what will happen:

Say, airspeed is suddently reduced due to wind shear. This is the only change, and AoA remains truly the same.
The moment balance remains the same, becuse lift of the wing and the tailplane changes proportionally. Thus no pitch change yet.
Lift will drop and will not sustain level flight anymore.
The aircraft will start accelerating down (while still in the same attitude!)
This acceleration has an effect of increasing AoA (due to the vertical component of speed). Notice that this angular effect depends on the absolute forward airspeed, which suggests that the effect will be lower at higher speeds.
Now the normal short-period static stability (the $c_m^{\alpha}$ or $c_m^{c_L}$ derivative) will cause a nose-down moment (if the system was stable) and will restore the original trimmed AoA.

This restoration effect normally happens much quicker than the speed change, especially for larger airplanes. So with a good approximation you could analyse just speed changes and assume AoA = const throughout the process. But physically this is not true, which, I guess, is the source of your confusion.

With the nose down, the aircraft will gain airspeed.
QED: the aircraft tries to negate the initial airspeed upset, which is the definition of static stability.
However, what happens next can be interesting. Due to high inertia, the aircraft will inevitably gain more airspeed than needed, and the process will reverse. Oscillations will ensue, and this is what is usually called "phugoid motion". The dynamics of this motion depends on several factors (absolute speed was already mentioned). Typically, this motion is much less dampened than the short-period AoA stability, and can even be dynamically unstable (i.e. have increasing amplitude). But this is not necessarily a problem: because the motion is slow, pilots can control it.

If you want derivatives, they have already been mentioned. Omitting derivation and with some simplifications, they combine to
$$\sigma_V = c_m^{c_L} - \frac{V}{2c_L} c_m^V$$
This $\sigma_V$ is the measure of static stability by airspeed. It is negative for stable aircraft. As can be seen, aircraft without direct speed-moment response ($c_m^V \approx 0$ - think gliders, for example) will have speed stability fully determined by the short-period AoA stability.

Re "Specifically, you seem to think that the tail load is zero in the trimmed condition. Not true. It could be true only for aircraft with neutral static longitudinal stability"-- hasn't it been established elsewhere on ASE that the tail load for an aircraft with positive static stability may be either uplifting, downlifting, or neutral?

From the static stability perspective, yes.  That is basically just the airframe weathervaning in the vertical plane due to a gust.  The AOA/speed seeking restorative tendency is TRIM and trim is a function of the opposing pitching moments, of ND moment from the wing/CG balanced by tail downforce, reaching equilibrium at trim AOA, whatever it's been set to with the trim control.  The airplane pitches when displaced from trim AOA/speed because the trim forces were moved out of balance, and the interaction of the out of balance moments creates pitch movements as it seeks to regain balance.

Inertia and tail volume are also contributory, adding variables.  A very large tail torque can contribute to AOA overshoot, too small and we get airspeed overshoot.  Zeus bullet 7 is a gem.

@quietflyer, generally, you are right, although for stable aircraft this (zero tail load) will happen only at a certain AoA. Only for neutral stability this can be achieved throughout the whole range of AoA. (It then raises a question why would we need such tail at all. Perhaps for damping, which Robert hints at). I implied this general case. I should have worded more carefully, but I always fear excess of side details when explaining basics: I have a propensity for this and it tends to confuse people...

With a large speed (lift) change, more tail volume can make the oscillation worse by overcorrecting, too small a tail does not correct fast enough.  Tails must be "right sized", but also large enough for directional stability and to assist stall recovery.  @quiet flyer is right that a plane can be staticly stable with an uplifting tail (all the way out to delta tail/straight canard), but the torques around CG must make the nose rise when it goes faster than trim and vice versa.

@Robert, what you are saying re tail volume is again a question of the [short-period] AoA stability rather than of the phugoid mode. It will afftect it indirectly as I explained, but without ensuring AoA stability first (by choosing the correct tail amongst other things), there is (normally) no point talking about phugoid motion.

@Zeus what I was proposing is short term is AOA overdone, and phugoid not enough  (allows speed to vary too much). In this way I believe they are related, sort of like the pitch version of Dutch Roll.  Anyways, as always, learning.  Thanks.

So that all brings us to:

alpha remains mostly constant

The old constant alpha variable speed phugoid caused by pitch inertia.

speed increases, which increases downward lift on the tail,
pitching the nose up.

With a staticly stable aircraft, the $torque$ of the tail (around the center of gravity) reacts more strongly than the $torque$ of the wing to a change in lift caused by a deviation from trim speed.  This means, as speed increases, the tail "pitches the nose up".  This is an increase in Angle of Attack, resulting in greater lift and a curving flight upwards.
As the gravity vector swings behind the plane (aircraft reference), it works with drag to slow the plane down.  This process repeats as the plane "seeks" its trim speed.
There is a very interesting interplay between changing speed (increases lift by its square) and changing AOA of the wing.  Phugoid is observed in very "slick" gliders, which have very low drag.
In the event of loss of headwind, lift drops, the plane sinks, AOA of wing increases, wing/tail torque imbalance lowers nose.  But the glider rapidly gains speed, resulting in a lift increase (and overshoot).  A draggier plane might be more "in balance" between AOA change and speed change.
This is where one might look to better understand longer period phugoid motion.  With weak tail downforce trim (such as when weight is moved back), the plane is more susceptible to speed change/lift differential oscillation.  Combined with very low drag, this would increase the chances of phugoid motion.  The "short period" motion may be other side of the spectrum, caused by excessive trim/weight forward causing a more poorly damped variation in AOA.

No, the torque (moment) doesn't change directly with the speed (for the simple wing-tail system). It only changes indirectly because of the AoA changes. In phugoid motion, all the lift changes are caused by speed changes (to the first approximation).

@Zeus The tail "reaction" is caused by changes in relative wind caused by a change in lift due to change in airspeed.  Thanks Zeus, absolutely correct that the moments initially are the same when airspeed changes.  Interesting it is broken into "short period" and "phugoid".  What actually happens may be somewhere in between.

The difference between short and long period motion just "happens" for real airplanes (esp. larger ones; it is not valid for paper airplanes). It turns out that the angular motion (AoA/pitch variation) is much quicker than linear motion (speed/lift/altitude), and this is handy for analysis: we can focus on them separately. It's not "in between", its both. But as a result, we can consider AoA=const throughout phugoid motion if we are only interested in its dynamics rather than physics.


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Swept back wing theory
I'm a beginner in airfoil design. I just came across swept-back wings. There is a point that I couldn't understand that theory.
Why is the only perpendicular component of the airflow to the wing(component parallel to chord line) affects the shockwave formation? Why doesn't the total airflow vector(spanwise +chorwise vector component) contribute towards shockwave?
Why wouldn't the shockwave form with the total airflow vector in the swept wing?
I have attached the image to show my point better

Imagine a wing that is infinitely long, and stretches from east to west. It is moving north at 400 mph. Obiously, it does not produce a shockwave, as it is moving subsonically.
Now, imagine a copy of this wing  that is moving to the east at 800 mph (along its length) while still moving north at 400 mph. Other than skin effects, this wing is aerodynamically identical to the first, so it does not have shockwaves, yet it is very supersonic.
Of course, a real swept wing is not infintely long, and so there are shockwaves at nose and tip of the wing, but mid span it really does behave similarly to a wing moving much more slowly perpendicular to its leading edge.

I'm talking about swept back wing. And I couldn't understand your answer. Could you brief your answer?

I propose you first read this answer.
The flow over the wing is accelerated along the pressure gradient (or orthogonal to the isobars), and this gradient is swept with the wing. Therefore, acceleration and deceleration is orthogonal to the direction of wingspan. The spanwise speed component remains unaffected.

source: Stack Exchange

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What should be the ideal distance between wings in a multi-winged air plane?
What is multi-wing or multi-plane:
For those who may not know, multi-wing air planes or multi-planes are aircraft that exceed the three sets of wings.

The idea is while increasing the number of wings, you increase the lift at the cost of drag.
This is also true for multi-element wings, or flaps.

The question:
What should be the minimum distance between multiple wings in a multi-plane configuration in order to avoid the airflow generated by the wings to interfere with one another?
Of course, I imagine that this depends on some factors, such as wing dimensions, speed and turbulence in the air.
But taking into consideration something like a the wings of this glider under speeds around 100 kph, what would be the minium distance between wings in that stair-like configuration?
Or the most practical and/or efficient way of doing that is putting the wings one under another like in a conventional biplane?

A multiplane might work fine at a specific AoA, but be a total (airflow) mess at another.  If AoA is within a narrow range, multiplanes can provide greater lifting force.  Jet engine compressors and fans are multiplanes.  Not as efficient as monoplanes, but with plenty of power, very effective.

@RobertDiGiovanni What is "AoA"?

AoA is Angle of Attack

The glider in your link has only one wing  How many wings do you have in mind? A couple or a dozen? Flaps and bi (or tri or multi) planes  work in a very different way.

@sophit at least a dozen

Something like this? What do you want to achieve?

@sophit kinda like that but more like in a "stair" configuration. I want to achieve a lot of lift to lift heavy stuff without the giant wings.

Rule of thumb is at least 1 chord length of spacing.  "Staggering" the wings can also be helpful.  The aviator/inventor pictured in your question seems to be off to a good start.  Multiple wings can be cross-braced and can very made very strong, which was an early advantage for biplanes.
However, once more powerful engines made higher speeds available,
advances in building materials and internal truss structured spars made the less draggy monoplane design the future of aviation.
The relationship of thrust and drag is best understood using a simple rock and parachute model.  Double weight (propulsive force), and speed increases by ... $\sqrt{2}$, or only around 40%.  This is why streamlining becomes imperative as speeds increase.  More wings simply become more weight and more drag.
The good news is that Lift also increases by velocity squared, and higher velocity gets you there faster (especially into a head wind).  So the strategy becomes less wing area and more speed.
Lifting factors are explained in the Lift equation

Lift = air density × Area × Coefficient × Velocity$^2$

Reynolds number  has a strong effect on lift to drag ratios.  Effects of airspeed on L/D ratios for various airfoils can be studied here.

Reynolds Number = Velocity×Chord/Kinematic Viscosity of air

Faster speeds give higher Reynolds numbers for a given wing.  Effects on L/D become very significant for an Re > 1 million vs 20,000.
Multiwings live on to extend the low speed range of STOL aircraft and very heavy airliners.  The ability to retract these devices greatly lowers drag at higher speeds.
One notable exception to all this is the eagle/vulture design, which favors very slow flight, minimal rate of descent, and the ability to ride even the slightest of updrafts.  Here, the thought of a modern glider sprouting multiple heavily cambered wing tips, slats and flaps in order to slow way down and ride the thermal seems a possibility.  Again, extra weight of these devices would have to be considered.  It would be fun to try.

I've seen a photo of an experimental plane with eagle/buteo/vulture-wingtip-like wing extensions, like an array of winglets aimed mostly horizontally but tilted somewhat upwards. The ones near the TE of tip were angled up higher than the ones further back.  At least 3 or 4 total.   I think the wing was a rectangular wing of not too extreme aspect ratio, and my recollection is that it was a low-wing plane resembling a crop-duster, and this may have been the intended use (not sure).  Would make a nice addition to this answer if a photo could be found but I can't recall where I saw it.

Of course, what the actual birds do, is open up those wingtip slots maximally when flying slow-- at increasingly higher airspeeds the whole wing is progressively more and more folded but the tip area is basically the first to fold, closing the slots--

@quietflyer back at R/C groups on of my early mentors marveled at the ability of old-time undercambered gliders to do well in competition with very light updrafts.

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What is the purpose of wing stiffeners?
There are very few articles as to what wing stiffeners are on the internet, so my question is what are wing stiffeners?

When you have a thin sheet that is put under a shearing load like a spar web as shown, the sheet, being flexible, has very little resistance to bending under compression out of its surface plane (or buckling) at which point its ability to resist the shear force is lost.  The stiffener just gives the flat web some resistance to compression buckling by making it less flexible; "stiffening" it in other words.


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How can an airplane roll, yaw and pitch up and down without control surfaces?
This is a Sprat 103 ultralight, it has no ailerons, and no ruddervator (on its v-tail). How then does it fly controllably?

Early aircraft had no flight control surfaces either. They achieved control by flexing of the wing structure, much like birds do.

The wing is split to two halves, each half is mounted on its own axis. By changing each half's angle of incidence (and thus angle of attack) changes the amount of lift generated by the relevant side.
See this video of a scale model.
A scheme of the mechanism

That gives roll.  Moving the two halves together changes the incidence angle (aka decalage) between wing and stabilizer, giving pitch control -- and the vertical fin is sized to avoid excessive side slip and make the craft spirally stable, like a free flight model airplane.


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How does drag change with Reynolds number?
If you want to scale down an airframe, i.e., for wind tunnel testing, you have to adjust for the Reynolds number. I want to ask if in case I forgot to adjust the Reynolds number, would the drag on the airframe be greater or smaller after scaling down than what could be predicted from change in frontal area?

Possible duplicate of How does Reynolds Number affect skin friction drag?

Recall the definition of the Reynolds number:

Where ρ is the density, v is the velocity, μ is the viscosity of air and l is the characteristic length.
The characteristic length is typically the chord length of an airfoil but can also mean the overall length of the airframe. When we scale down a model, we are decreasing l.
Decreasing l while holding all other variables constant (which I assume is what you meant about "forgetting to adjust the Reynolds number"), has the net effect of lowering your Reynolds number.
Looking at airfoil graphs for example, we can see the relationship between the Reynolds number and the drag coefficient of the body:

It is obvious from the graph that the drag coefficient increases when the Reynolds number is lower (higher curve on the graph)
Image source: Airfoil tools


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How would the wave drag of an elongated Sears–Haack body change with a cylindrical mid section?
Would a Sears–Haack body have the same wave drag as the same body with cylinder of the same diameter placed in the middle, between two halves of the Sears–Haack?

No it will not have the same wave drag.
The low wave-drag of a Sears-Haack-Body is achieved by the smoothness of the body. This means kinks are not allowed. The picture below shows the Mach-Number distribution around a Sears-Haack-Body (taken from YAMAZAKI 2014). Notice the gradual change of the Mach number along the body surface .

The following illustration shows the proposed elongation of the body by inserting a cylinder.

By adding a cylinder the curvature of the body does not change gradually (smoothly) but stops - stays constant - and starts again.
This will trigger a Prandtl-Meyer-expansion (example below, taken from here) which will distort the smooth flow around the body and thereby increase the wavedrag.

Wave drag of a body of revolution depends on volume, so increasing the volume of an already minimum wave drag body will increase its wave drag. When you lengthen the center section you increase the volume of the body and therefore the drag, and since the resulting shape is not Sears-Haack the drag is even higher.
See Computing Supersonic Volume Wave Drag halfway down the page.

Actually, according to the link drag is proportional to v^2/l^4. Since im only making the body longer volume and length change by the same ratio, and as the length is raised to higher power the drag will actually decrease, correct?

@FrancisL. Look farther down the page. What it means is that wave drag depends on volume, and a Sears-Haack body gives the lowest drag for a given volume. When you lengthen the center section you increase the volume of the body and therefore the drag, and since the resulting shape is not Sears-Haack the drag is even higher.


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Do FAA regulations keep aircraft away from 18,000 feet MSL to avoid conflict with FL190?
In the US, is there any FAA Regulation that keeps all aircraft away from 18,000 feet MSL… Since theoretically, FL190 could be lower than 18,000 feet MSL, and create all kinds of confusion with keeping safe aircraft separation?

Can you perhaps make your question more specific? As it is right now, the title itself isn't a question and in the body you seem to ask two different things: a) are there regulations that prohibit flying at 18,000 MSL, and b) how is separation provided close to the transition level/altitude? It's not clear to me at least what your question really is.

It's not the altitude that is adjusted, but the flight level. Aircraft are allowed to fly at 17,500 feet MSL. Per the rules provided in other answers, this would be VFR flight to the north or east. an IFR aircraft could be flying at FL180, and this would provide 500 feet of separation.
However, 14 CFR 91.121 determines the lowest usable flight level. If the altimeter setting is 29.92 or higher, FL180 provides at least 500 feet separation from the next VFR altitude, and at least 1000 feet from the next IFR altitude. However, if the altimeter setting is lower, then the lowest usable flight level is higher than FL180.
Current altimeter setting   Lowest usable flight level
29.92 (or higher)           180
29.91 through 29.42         185
29.41 through 28.92         190
28.91 through 28.42         195
28.41 through 27.92         200
27.91 through 27.42         205
27.41 through 26.92         210

Thank you, Fooot.  THIS is what I was looking for.  At certain times, what I was theorizing is, in fact, a valid concern, and 14 CFR 91.121 speaks to it.

That's where the transition level comes into play. When climbing through 18,000 feet MSL (transition altitude), the altimeter will be set to the standard pressure of 29.92 in.Hg or 1013 hPa. However, when descending, the local QNH will not be set at the TA, but the transition level (TL). The transition level is dynamic and depends on the local air pressure.
Following a table on how to determine transition level from Wikipedia:

Source: Wikipedia
The switch to local QNH happens, when passing the determined TL.
The lowest flight level one can fly is TL + 500 ft but many countries also add a buffer to TA.
So that's how aircraft can fly at both TA and TL while still having sufficient vertical seperation. (Transition layer + buffer)

Thank you, Noah. I thought I made it clear within the context of my questions that I’m well aware 18,000 feet MSL is the transition to pressure altitude. While your answer is quite accurate, I don’t feel it addresses my question. The fact is that some VFR Aircraft may, for example, wish to fly eastbound at 17,500 feet MSL. If at the same time, a slightly higher IFR Aircraft were at flight level 190, absolute separation could be anywhere from 2,500 Feet to Zero, or even inverted ... with the aircraft at FL190 flying below the aircraft at 17,500 MSL.

Once in the Class A airspace all a/c are IFR. Based on the QNH, ATC won't assign a flight level that puts an aircraft at an MSL altitude that could conflict with traffic at or below 18000 feet. With a QNH of 950mbar, the lowest FL ATC will assign is FL210. Flight levels below that are not usable for the reason you state.

Many countries have a variable transition level ensuring the transition layer will always be at least 1000 ft (typically between 1000 and 1499 ft), meaning that the TRL and TA are vertically separated. For some reason, as with many other things, the USA feels the need to be different, in this case with a fixed TA of 18000ft and a fixed TRL of FL180.

To clarify; air traffic procedures exist to avoid the problem. They can be found here.


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Why are push-propellers so rare, yet they are still around?
Jet engines are by their very nature push-engines, however, most propeller airplanes use pull-engines. Is there an inherent advantage to using pull-propellers except for the increased airflow over the fuselage and tail (with its rudder and elevator)?
Twins generally have their engines on the wings, and the tail is no longer directly behind it, does that mean the choice of a pull-engine is not as advantageous?
If there isn't an inherent disadvantage, why are pusher configurations so rare? If there is one, why do they exist at all? Disregarding designs where the choice is obvious, like powered parachutes where you simply don't want a propeller in your face.
The Convair B36 is one notable multi-engine aircraft with engines in pusher configuration, as is the Piaggio Avanti. And the Cessna Skymaster is a push/pull configuration (If you get a multi-engine rating in a Skymaster, your ticket will be limited to multi-engine aircraft with in-line thrust). Single engine aircraft are even more uncommon, and pretty much all I could find except the Lake Buccaneer are all kit-planes (e.g. Velocity, Rutan), ultralights (Quad City), military, or experimental.

You can make a "puller" jet, but no one does. Look at the airflow of the PT-6 turboprop engine. Notice the intake below the exhaust, the back-to-front combustion path and the exhaust pointing backward almost at the front of the engine. Remove the prop and gearbox, optimize the engine for pure-jet, and you have a puller jet engine. Should be obvious that the other designs are more efficient. Early rockets were built in the puller configuration, one production jet close to a puller would be the Pegasus, used in the Harrier - it has no rear-facing exhaust, using nozzles to vector the thrust.

@Federico you were right. Note that turbojet engines are not real push engines. The force is not generated at the rear of the engines. Turbofans are very much puller engines with most of the thrust generated by the compressor. Pure jet engines like in the Tornado are perhaps a bit more push than pull, but still it has a lot of pull in it. Take a look at the strain in the main shaft. For pull props, the shaft is tension loaded, for push props the shaft is compression loaded. For all jet engines, the shaft is tension loaded; the compressor pulls the aircraft ahead.

@Federico no thrust is generated in the turbine, on the contrary: the turbine extracts energy from the internal airflow to drive the compressor thereby creating drag. The combustion chamber provides most of thrust in a pure jet engine.

The Caspian Sea Monster was a jet puller so it could deflect jet exhaust downwards under the wings for extra lift.

If the jet engines are in the wings like most large commercial airliners, then it's actually a push/pull. The rear half of the cabin is pulled along, and the front half of the cabin is pushed along.

For piston engine airplanes you need the engine in front for to move centre of gravity in the front for stability. In this case if you need a pusher, the transmission shaft need to go through the carbin. Or you go with rear engine but the engine has to be light to make it easier to design the counter weight.

Modern jet engines are actually "turbofan" engines, in which most of the actual thrust is produced by the large bypass fan at the front of the engine.  The fan is in turn powered by the turbine at the core, which does produce some thrust of its own at the back.

@user3528438 -- re comment above -- you would be mistaken there.  Many Rutan designs used pusher engines of normal design-- and no transmission shaft through the cabin.

There are many disadvantages, they seem to outweigh the advantages.
Here are two:

A pusher prop is working in a disturbed airflow, causing increased vibration and noise
If the propeller is fitted behind a wing, each propeller blade is passing through the separated boundary flow twice each rotation. These cycles create additional noise and lower the efficiency of the propeller. The vibration makes the propeller blades more susceptible to metal fatigue.
Propeller clearance at takeoff
Due to the pitch up at take-off, the propeller gets close to the ground. Therefore the diameter needs to be reduced (loss of efficiency) or the landing gear struts need to be made longer (added weight). Since the propeller is behind the landing gear,  it is susceptible to debris kicked up from the gear, increasing the need for added blade protection (increased weight, loss of efficiency)

Wikipedia has a list of additional disadvantages.

Edit:
Your statement that

Jet engines are by their very nature push-engines

is not entirely true.
In turbofan engines, most of the thrust is generated by the fan and compressor stages. Even in a pure jet engine, a lot of thrust is generated by the compressor. Therefore the shaft of a jet engine is tension loaded, just like a propeller shaft in puller configuration.

Source: Rolls Royce - The Jet Engine

I noticed this list as well. So why are there pushers at all?

@falstro you can have an unobstructed view out front, they also are quieter since the engine is at the back. they may be safer as well, since they are difficult to stall if they have a canard configuration: http://www.youtube.com/watch?v=H50zFi11OMU (just to name a few)

@falstro Just above that list is a list of advantages. That are the reasons people built pushers. And once they are built, they are around for a while. Although you don't see them often.

The interference with the wing wake is a serious issue, but only when the prop axis is in the plane of the wing. Once it is lifted above or lowered below, the prop blade will slice through it gradually than hitting it over the whole span, making noise and dynamic loading tolerable. Please take look at the Piaggio Avanti: This is an excellent design.

@DeltaLima I'm not convinced the turbofan or jet engines are either push or pull just because of loading on the shaft; helicopters often have turbine engines, yet it's the fact that the rotor is on top pulling it up that makes it a puller configuration, isn't it? Having the rotor on the bottom, would make it a pusher, which of course would be insane. Not impossible though, space shuttles would be an example of a vertical take-off craft in pusher configuration.

@PeterKämpf looking at the Avanti is ok, but have you ever listened to it? I really like it but it's not exactly silent. It is the example of a loud pusher prop design.

Also, pushers allow higher tail angles without having to worry about airflow seperation, and they decrease scrubbing drag and increase laminar flow because there is no propwash on the aircraft.

@falstro: That diagram does not show the loading factor on the shaft. It shows the thrust and drag contributions of different parts of the engine. The compressor stages generate most of the thrust compared to the hot gas expansion at the tail. So a jet engine is more of a propeller engine where the prop is ducted rather than an air-breathing turbocharged rocket engine.

what is the 5587? Is it drag of the engine or something?

The pusher design is more efficient, because the suction forward of the prop reduces flow separation, and the accelerated flow behind it is not streaming around the fuselage (or wing), where it would create additional friction drag. In case of the Do-335 (see picture below, source), the single-engine top speed was 30 km/h higher with the rear engine running than with the front engine (both were DB-603s with identical power rating).

On the other hand, the puller prop will help to maneuver the plane on the ground (this is a big benefit for taildraggers - note how many two-engined, taildragger airplanes have an H-tail (two rudders as endplates of the stabilizer). They were placed in the prop wake and this gave much better directional control at low speed on the ground. Also, the prop wash helps to increase the lift from flaps.
The main disadvantage for a single-engined aircraft, the reduced tail clearance, has already been mentioned. If you cannot really rotate, takeoffs and landings are high-speed affairs. More disadvantages are the pilot risking to hit the propeller in an emergency bail-out and increased stability of the aircraft. When developing the Saab J 21 (pictured below), Saab also developed their own ejection seats to allow safe egress.
A pusher propeller influences stability much as an additional tail, but without control surfaces.  Especially for a fighter aircraft, this is the opposite of what you want. THAT is why almost all high-powered, singe-engined aircraft have their propeller in the front: Maneuverability!
The stabilizing effect increases in proportion to the propeller surface area and the thrust, of course. Since a regular airplane needs to have basic stability with the engine running at idle, any additional stability change due to propeller placement comes on top. At full power and with the long lever arm of a single pusher prop on a central fuselage (think LearAvia Learfan), the aircraft becomes stiff as a brick. A two-boom layout (think Saab J 21, pictured below, source) is better, but creates additional friction and interference drag, so the advantage of the pusher arrangement is reduced. Note, however, that studies of a tractor engine variant (Saab J 23) showed inferior performance to the pusher design.

If you want hard data on that: There is an old NACA report (NACA TN 2585) on this by John L. Crigler and Jean Gilman, called Propellers in Pitch and Yaw.

stability depends on overall design, not just engine placement. If you use a twin boom tail like the D.XXIII or Cessna Skymaster used, that problem largely goes away. But yes, it can be a factor. And of course if you use such a design, you're more than halfway towards a design with enough space to put both engines in the wings, by extending the twin booms forward to become engine pods, like in the P.38. The push/pull design however makes for far better engine-out handling, as there's no off axis thrust vector.

I've read that the stabilizing effect is so strong that pusher-prop driven flying wings didn't need vertical tail planes/rudders e.g. the Northrup flying wings. When they converted to jets, they had to add them.

Fun fact: Both of the P-38s engines were critical engines, because the other way, the propwash made the nose wander a little (not good for aiming your guns!)

@ptgflyer I always knew the P-38 had both engines critical, but I never knew why until now.  Thanks!

"It stabilizes the aircraft, much as an additional tail, but without control surfaces."-- it's hard to imagine that this is as much of a disadvantage as you imply, for fighter aircraft.  Consider the enormous size of the vertical fin of the Blackburn Firebrand.  How does keeping the fuselage streamlined to the airflow really restrict the available turn rate?

@quietflyer: It's about pitch and yaw damping. If you don't believe it, it will be best to build a model and test it yourself. Besides, the Firebrand needed the big tail to counter the gyro effects of its radial engine.

@PeterKampf -- if you were testing the max available roll rate available WITHOUT deflecting the rudder-- which is sometimes how these tests were done in the WW2 era, I've seen a write-up-- wouldn't the roll rate be higher with a big vertical fin to reduce adverse yaw?

@quietflyer: Why should a fighter roll? To change course, and for that it needs to pull gs right after rolling. This pitching motion becomes more sluggish with a tail mounted prop and requires a bigger horizontal. In order to have sufficient pitch stability for approach and landing (when the engine is idling), the pitch stability with full power becomes crippling for air combat. Roll isn't affected, but it is not all what counts in air combat.

So the fighter is unable to pull the desired G's even with the stick well aft?  Or there is a time lag involved in getting the desired angle-of-attack / G-load established?  I can see your point especially in terms of the pitch axis, not sure I'm convinced in terms of the yaw axis.  Thanks for the note though.

@quietflyer: Both: The response is slower and the maximum rate diminished. And the stick forces are higher for the same pitch rate / load factor.

@PeterKämpf Do you have a source for the 30 km/h difference? I fully believe it, but would be interested to see some performance data of the aircraft. In the flight handbook I found online there was no info about it

@Gypaets William Green, Warplanes of the Third Reich.

@PeterKämpf It is so nice to see that after producing over a staggering 2100 answers, you are still giving your third-ever answer on this site a little bit extra love and attention to make it even better. Makes me smile!

Having been reading up on 1930s and 1940s aircraft, I can think of a few major problems that dogged the design back then.
Pushers had advantages of pushing the wings and control surfaces through undisrupted air and that gave enough advantages that the a lot of pusher designs were put forward. A pusher would have superior cruising speed and better wing loading because it flew through undisturbed air.
The most successful was probably the XB-42 mixmaster which would have been deployed had the war gone on. (As it was, it it eventually deployed as America's first jet bomber.)
The B-36 used pushers precisely to give its wings a clear airflow. It was actually designed in 1942.
But they had big tradeoffs.
A major problem with pusher designs of that era was cooling the engine on the ground.
With a tractor (puller) you have this giant fan running blowing air over the plane and the engine (air cooled radial) and/or radiator (water cooled.) With a pusher, the prop just blew air behind the plane, leaving a powerful engine with nothing but passive cooling.
During a crash, a tractor engine in front served as a crumple zone for the rest of the plane, particularly the cockpit and could plow through obstacles making the plane come to a halt slower. Conversely, with a pusher, the engine was behind the pilot and not only didn't offer any protection but tended to tear loose and ram through the rest of the fuselage like a pile driver.
The pilot could not see and therefore visually inspected the engine e.g. if you start loosing oil in a tractor, you know right away. If the engine is behind you, you might not notice until you see the gauge.
Center of thrust was more difficult to balance and that made center of gravity tricky as well.
Still, with all these disadvantages, XB-42 and B-36 showed that the advantages could be come out on top with enough good engineering. And in the end, jets work at least half by pushing. If jets had be delayed a few years or the war had started back in the mid-30s, we probably would have seen more pusher designs in service.
With jets, the military need for pushers disappeared and in civilian aircraft, there is seldom much need for an increase in performance that justifies fighting all the tradeoffs.

The B-36, interestingly, had the opposite of the cooling problem; the engines it used were designed to be used in a tractor configuration, where the air intakes would be located behind the cylinders, which would warm the air coming into the intakes.  With the B-36, it was the other way around, with the intakes out in front with no heating whatsoever.  Consequently, the B-36 was very prone to engine failures caused by carburetor icing.

Why are push-propellers so rare

Other answers have covered this.

yet they are still around?

A front mounted propeller limits the field of view of the pilot or payload.
Consider a surveillance drone or UAV. There's no pilot whose view would be obscured but there is almost certainly a great advantage for the radar, optical and other forward-looking systems.
Surveillance missions can benefit from a relatively slow-moving, low-altitude platform. This favours the use of propellors rather than jets and the pusher arrangement can be beneficial as noted above.
  MQ-9 Reaper
 Elbit Hermes

Sagem Sperwer

Just some additional info and thoughts on this old thread.  There are still a few successful pusher aircraft, even a few being built today.  The Piaggio Avanti is quite successful and very fast, high cruise is about 400 knots, and there is the above mentioned CBA-123 in development now.  Like all engineering design, pusher aircraft are a compromise.  You typically lose some propellor efficiency, from the disturbed airflow, but gain some aerodynamic efficiency from the slower airflow over the wings and fuselage.  Prop clearance can be an issue, as is FOD of the props from debris picked up by the wheels being thrown into the props.
Noise is a mixed bag.  The internal noise is often lower, because the props are not blowing onto the windscreen or fuselage, but external noise can be an issue due to the disturbed flow to the props. The Piaggio Avanti is supposedly one of quietest cabins flying but it is quite noisy when it flies by.
I had a  pressurized skymaster for a number of years, and I really liked it, but it was very noisy both inside and out.  But the skymaster was very quiet inside with just the rear engine running so it wasn't the pusher prop that was the issue.  Performance was slightly better on the rear engine alone.  The "wives tales" about taking off on one engine are just that.  Like all piston twins the performance on one engine was marginal, and there is no way that a proficient pilot wouldn't notice an engine failure.  I did fly the plane on one engine a number of times, both during simulated engine failures on takeoff and one actual engine failure in flight, and while the handling was docile the performance was unimpressive (though a P-337 has somewhat better single engine climb rate than most piston twins).
I did many rough field operations in my Skymaster and was lucky to never have damaged the rear prop but there is little doubt it could be an issue.  I have not heard of issues of prop damage on Piaggio Avantis but they aren't likely to be used on unpaved runways. So, like all airplanes, the designer needs to consider a host of issues, including the intended use of the aircraft, before deciding whether to use a tractor or pusher configuration.  I suspect that much of the reason for the lack of pushers is the inherently conservative attitude of aircraft designers, not the inherent issues with pusher design.

I did some reading up on the Avanti - a very elegant design, from an engineering point of view as well as aesthetically.  An advantage of the pusher-turboprop layout is apparently that the prop blades don't need de-icers. because the exhaust stream does it inherently.

For single engine aircraft at least, having a pusher prop makes egress in flight (a.k.a. bailing out or ejecting) much more dangerous. During WW1 I believe (when pusher props were far more common) more than a few pilots were seriously injured or killed by their props when having to jump out of a burning aircraft.
For this reason the Germans fitted (or planned to) an ejection seat in their Do.335 push/pull design, and Fokker planned to do the same in his D.XXIII.
The B-36 was large enough that pusher props would be far away from the fuselage, getting rid of that problem.
An advantage I can think of of having the props in the front that stems from the increased airflow is that you get free extra cooling of the engines.
Having the engine in a single engine design in the rear with a pusher prop, you end up with a more complex air inlet and tail design than simply letting the engine suck in air that it bites into already. You need ducts and stuff, adding weight and complexity.

The Do-335 did have a pneumatic ejection seat, and additionally the upper fin and the propeller were blown off by an explosive charge when the seat was activated.

Does a pusher propeller run in disturbed airflow, producing more vibration and noise? Yes.
Does a pusher engine have more cooling issues? Yes if it is a piston engine (turboprops are cooled by their own internal airflow).
However there is one thing in favor of the pusher configuration which needs a better evaluation: A pusher propeller picks up airflow that is already decelerated by the wing and fuselage to generate thrust, while a puller propeller picks up undisturbed air to blow it onto the fuselage and wings, which in turn decelerate the airflow. This difference probably means that a pusher engine may be somehow more efficient at higher speeds. A twin pusher, shown in the following photo, was developed here in my country in the 80's called CBA-123 (two prototypes were built) and the experience with that airplane points in this direction.

They should have made one engine push and one pull to solve it once and for all.

One category of aircraft that almost universally uses pusher props is the autogyro.  Typically the engine is mounted in the centre of the craft, with the cockpit immediately forward, the rotor above, and the tail somehow mounted behind the prop - which requires an often spindly-looking strut wending its way past the propeller disc.
The same is true of some other types of ultralight aircraft, with a small engine often being strapped directly to the pilot's back, or possibly the back of his seat.  In most cases that's the only place it can be mounted without starting to build an actual fuselage.
In both cases, there is some sort of guard which would strike the ground before the propeller itself if sufficient pitch were attained, and the pitch required for this is much greater than most aircraft need at takeoff.  A propeller strike is therefore unlikely.
The major advantage of a puller-prop configuration, from an aerodynamic point of view, is that it adds forward airflow over the wings and tail in the most crucial phases of flight, when the aircraft as a whole is going rather slowly.  Before the jet era, long tarmac runways were somewhat uncommon, so being able to take off from a short and/or rough field was advantageous.  For fighters, the extra few seconds of controllability while manoeuvring in the vertical could easily mean the difference between victory and death.
This was of particular importance for carrier-based aircraft, which during WW2 did not typically have the benefit of a catapult to accelerate them off the deck.  The carrier could generate a significant headwind to help out, by steaming upwind as fast as it could, but a heavily loaded fighter or light bomber still needed as much of a safety margin as it could get.  The B-25s used in the Doolittle Raid were a truly exceptional case; the crews needed special training in order to take off from something as small as a carrier, and one of them only managed it by floating in the "ground effect" just above the waves after leaving the deck.  If they'd been pushers, they probably wouldn't have made it.
Also relevant to carriers is the ordeal of landing on a deck.  The universal configuration for any non-VTOL carrier aircraft is a hook, coupled to a heavy-duty shock absorber in the tail, which engages with a series of heavy cables laid across the flight deck.  To make that work, the plane basically needs to fly itself into the ship in a nose-up attitude, with none of the flare you'd usually employ when landing on a proper airfield.  A rear-mounted propeller would routinely strike the deck in that case, becoming a major hazard to both the aircraft itself and the flight deck crew.
For these and several other reasons, the front-mounted propellor became conventional in military designs and, by extension, in civilian designs as well.  A pusher automatically seems unconventional in that context, and is associated with ultralight and otherwise unusual aircraft, such as autogyros.
The shift in CG associated with moving the engine to the back also encourages major changes in planform even with full-sized aircraft, up to and including canard configuration, which makes such an aircraft appear even more outlandish.  A canard aircraft also has different handling characteristics at the edge of its flight envelope, which may put prospective pilots off even further.  More recently, the three-surface layout (with both canard and conventional tail) has appeared, carrying at least theoretical advantages from both canard and conventional layouts.
Few aircraft designers are willing to stick their neck out into such an environment; that the Beechcraft Starship is widely regarded as a failure - Beechcraft actually destroyed many Starship airframes in an attempt to avoid the costs of supporting them - would tend to encourage further caution.  Burt Rutan (of Scaled Composites) is a notable exception to the rule; he has designed several successful canard and three-surface aircraft with pusher engines.
The Piaggio Avanti, however, is a rare example of an aircraft which takes the pusher's inherent advantages to their logical conclusion.  With a three-surface layout and very careful attention to aerodynamic details, it is a turboprop aircraft with very close to corporate-jet performance, yet with significantly lower operating costs and the ability to operate from shorter runways than a jet.  Being backed by Ferrari, including ferrying the Scuderia racing team around in an Avanti, can't have hurt its popularity.  Perhaps more remarkably, the only significant accident I can find involving an Avanti was the loss of an unmanned drone based on the type - not at all bad for an aircraft type that's been flying hundreds of examples for nearly 30 years.
So the pusher configuration has its own advantages, which most aircraft designers don't really know how to make the most of because jets and front-mounted props have been the standard for so long.  But the latter became dominant due to decisive advantages it held in the pre-jet era.

Can you provide more detail about the benefits of pushers for autogyros?

I'm not certain if there are any major aerodynamic advantages, but an autogyro needs the rotor directly over the CG, and with a clutch to the engine to spin it up for takeoff, so putting the engine right in the middle is natural.  From there it seems mechanically simpler to install a pusher prop, at least once you figure out how to mount the tail.  I've read that an autogyro needs a low centre of thrust to be maximally stable, which may also influence the usual design.

You could edit that into your answer. The question is asking what the benefits are and your answer doesn't address that much, just that autogyros use that configuration.

To be precise, the question asks about inherent advantages of puller props, and largely discounts cases where a pusher is the only possible configuration for the type.  I think you could build an autogyro with a puller prop, it just usually makes more sense to use a pusher.

Pusher props offer a real advantage to small amphibians like the Lake Buccaneer and the Republic See Bee where pilots need to be able to jump out and catch mooring buoys, lines or jetties without fear of striking a spinning prop blade.

It's also easier to keep the prop out of the water without placing it right over the cockpit.

One advantage of the tractor not yet mentioned is that it puts the engine weight in the front. We all know that model gliders need nose weight and the same applies to the real thing. using the engine makes for a more compact, lighter and simpler construction.
The issue of crash safety has also been mentioned. Funnily enough, the original Biggles books were technically a unique historical record by an ex-WWI pilot, Capt. W.E. Johns. Biggles Learns to Fly contains a graphic and shocking exposition on the relative merits of pusher and tractor in a crash.
At the start of WWI many British types by the RAF and Vickers were pushers. Whan newer types like the Sopwith 1 1/2-strutter came along, they proved far more compact and agile - and safer in a crash-landing. The pusher became obsolete overnight. It has remained a niche option ever since.

The same as in Do-335 happened with a twin engine Cessna, the Skymaster, in single engine operation, it was better with the rear engine. This Dornier-335  airplane looks more as a racer than the fighter-bomber it was designed for.

The Do-335 was designed as a bomber, which actually better fits its design with a large empty fuselage and widely located propellers to ensure maintaining airworthiness when damaged.

@Jihyun: Really? The concept started as a point-attack bomber but was quickly converted to a fighter. The fuselage was stuffed full of engines, tanks and a pilot (or two) and had no space for bombs.

Yes. It did start out as a bomber. I'm not sure what you're saying because modifications after initial design don't contrast my point.

Pushers were tried during WW2 without much success, and this was a time when just about anything was being tried.
The US had two pusher fighters under development during WW2, the Curtiss XP55, and the Northrop XP56. Neither became operational, both had stability issues, although those issues derived more from the radical designs rather than the pusher prop. The XP55 used a canard at the front for elevators and wingtip rudders, while the XP56 was effectively a flying wing with it's short, stubby fuselage.
The XP55's basic layout was later used in the Beechcraft Starship. Sadly, the Starship was not successful, costing more and flying slower than contemporary executive transports like the Cessna Citation and Lear, or even tractor turboprops like the Piper Cheyenne and Beech KingAir.
Japan developed one pusher fighter, the Kyushu J7W Shinden. Its pusher design reflects the intent to power it with a gas turbine engine, the pusher/tailless design to get around the problem of hot exhaust melting the tail. Like the American efforts, it never saw production, although the decaying industrial situation in Japan in 1945 was more responsible for it's failure to enter service.
In the case of the fighter designs, the pusher configuration was just one of several radical departures from conventional design, so it's hard to determine how much the pusher configuration contributed to the problems encountered. Note that the DO335 was a push-pull configuration, not a pure pusher.
What can be said with some degree of certainty is that pusher configurations, definitely in fighter aircraft, were not found to be beneficial when prop fighters were at their peak.

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text:
Should you wait to switch to the ground frequency until after crossing a runway?
I've never had the opportunity to fly into a controlled airport with parallel runways, so I've never actually faced the situation. But, in the interest of being ahead of it, when would it be prudent to switch to ground frequency with this clearance (assume I just landed)?

N12345, right on E, cross 27 R, contact ground .9

Should I call ground (or switch frequency, so I no longer hear the tower) before or after crossing 27 R? My by-the-book assumption would be immediately, as there's no "then" or "after crossing" or similar, but somehow that feels wrong. I suppose normally it doesn't really matter, but in the case of someone (which includes me, everyone makes mistakes) doing something wrong, where would they call me as I'm about to cross 27 R?
The reason I ask is that I saw exactly this case in a YouTube video quite a while ago (can't find it anymore though), where exactly this happened. The pilot was narrating for the video and stated he'd switch frequencies "now that he'd crossed the runway", at which point I started thinking about it.
Considering the situation, I would probably rather wait since there's an inherent delay when switching frequencies anyway, I might as well wait a bit longer, but knowing myself, if I hadn't thought about it ahead of time I'd probably switch right away. The only reason I can think of for saying it this way is if 27 R is not an active runway, but then again, reality does not always match what's in the book.

In reality, you're most likely to hear "cross 27R and contact Ground .9 on the other side". That phrasing removes the confusion when to switch & seems to be quite common.

Remember that active runways are always the jurisdiction of the local (tower) controller. Even if a ground controller tells you to cross a runway, they have to clear it with the local controller next to them. So generally you switch frequencies when you have finished crossing the runway unless told otherwise.
I can tell you this is exactly how it works in Boise, Idaho where there are the parallels 10/28 L/R.
However, if you are unsure of any instructions a controller gives you, the basic rule is: ask. The tower controller won't mind, and will typically appreciate the clarification. Just read back the crossing instruction as normal, then ask.

Right on E, cross 27 R, contact ground .9, N12345, and should I contact ground on the other side of 27R or immediately?

This is how I have always done it.  In general, unless specifically told otherwise, I would keep talking to the guy who gave you the clearance to cross until you are actually on the other side just in case a situation develops and they need to contact you in a time critical manner.

and make sure your tail is completely clear of the hold short line!

N12345, right on E, cross 27 R, contact ground .9

Do it in exactly the order they tell you. Right turn on E, cross the next runway, then stop and contact ground.
Have an airport diagram handy if you can, it will make a world of difference.
ETA: If they want you to switch to ground before crossing 27R, they'll probably just say this:
N12345, right on E, contact ground .9

In which case, you turn right on E, cross the hold position lines, stop, and switch to ground. The ground controllers will direct you the rest of the way, coordinating with tower controllers where crossing runways is required. In fact, in this case, they'll probably give you very explicit instructions not to cross any more runways until told to do so.

The thing is, if you'd for example gotten an enroute clearance, "climb and maintain 6000, contact duckburg center 123.4" you're not supposed to wait until you're at 6000, you're supposed to contact duckburg now, unless you actually get "climb and maintain 6000, reaching 6000 contact departure" or something similar.

Hmmm; good point. Any controllers here?

I'm not a controller, but I just about guarantee the second option would be accompanied with a "hold short 27R." The FAA is doing everything they can think of to eliminate runway incursions.

Very good point.

Different airports will have different procedures. In some instances, exiting a runway and joining a taxiway upon landing means you should immediately contact GC. In others, ATL for example, you shouldn’t ever change frequencies until instructed to do so. Reason for that is that in ATL local control owns jurisdiction over certain taxiways - B, N and to a lesser extent R and points south. Jumping frequencies here without instruction will lead to lost comms, and is a common issue among pilots not familiar.
As to the you-should-never-cross-an-active-on-ground point made earlier - not accurate. We routinely coordinate among each other to cross runways if an advantage can be gained by doing so. I probably cross 30 aircraft a week on ground control frequency.

Welcome to Av.SE - good to have more active ATC controllers on the forum!

You have been cleared by the tower to cross a runway so you should stay with the tower frequency until you've crossed it. The reason being that ground control is not authorized to issue clearances for runway crossings.
In theory, you should never cross a runway when on ground frequency because the tower is in charge of that.

The ground controller is however authorized to order me to stop, effectively canceling the crossing clearance, isn't he? Also, I've had numerous runway crossing clearances delivered on ground frequency, probably issued by the local, but still, your last statement doesn't add up, I shouldn't switch to tower just because I'm about to cross a runway, should I?

Crossing a runway while on ground frequency happens quite often. The ground controller simply authorizes it with the local controller first. You shouldn't switch frequencies while crossing unless told to do so.

What if the tower clears you to cross and for some reason has to stop you. If you're already on the ground frequency that could cause a problem. Personally, I wouldn't switch to ground until after crossing the runway.

The scenario I gave in the comment is when ground clears you to cross, not tower.

@BretCopeland You're right. It seems I wasn't wearing my reading glasses :-) However, in the situation described by the OP I would never switch to ground until after crossing the runway.

@PhilippeLeybaert for that scenario, I agree.

except at the largest airports, ground and tower sit next to each other, and coordinating

When you're given a clearance related to a runway, you need to stay with whoever gave you that clearance until you're clear of that runway--and only then do you follow an instruction to change frequency.
Why?  If Tower needs to cancel or amend your clearance, they will first try to reach you on the same frequency.  If it's an emergency, the several seconds it takes to realize you're not there and then several more seconds to have someone else call you on another frequency could mean the difference between you living and dying.
Imagine, for instance, that tower told you to cross the right runway, and then they realized the aircraft behind you was (incorrectly) lined up on the right instead of the left.  They're going to tell them to GO AROUND, of course, but they're also going to tell you to either STOP IMMEDIATELY or NO DELAY (i.e. full power), depending on exactly where you are.  You really don't want to miss that instruction, especially since someone who's dumb enough to line up with the wrong runway is probably also dumb enough to miss (or mess up) the go-around.

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text:
If a hot air source was added to the wingtips, would it help control the formation of wingtip vortexes?
Let's say that somehow, an engineer made an exhaust pipe that constantly sprayed hot air at the wingtips in order to help control the formation of less turbulent wingtip vortexes.
Would that have any meaningful effect?

The difference in pressure between the lower part of the wing (high pressure) and the top part of the wing (low pressure) allows the wing to fly, and also generates the wingtip vortex (if I didn't understood it incorrectly).
So, if you add hot air at the wingtips, it could probably allow the formation of wingtip vortex away from the wingtips... Or not.

Given that the air temp at cruise altitude is ~(-44C°), what would be the benefit of said "sprayed" air being "hot"?

How according to you would hot air reduce wingtip vortices?

I edited the text to add: "The difference in pressure between the lower part of the wing (high pressure) and the top part of the wing (low pressure) allows the wing to fly, and also generates the wingtip vortex (if I didn't understood it incorrectly)."

Some designs have propellers there, others winglets (including wing tip washout) or fuel tanks.  But the best way to handle it is lower AoA and greater speed.  Not unlike a power boat going fast enough that it's turbulent wake never catches up .

By what mechanism do you think that this might possibly work?  Is this simply a random "what if" idea with nothing to suggest that it would in fact work, or is there some reason that you think that it might?

@RalphJ FWIW, something similar to a bleed air lines from an engine's exhaust (just aft of the last power turbine disc) could be a logical source of hot air, but running that hot air through the wings (where the fuel tanks are) may not pass the 'sniff test' during engineering design review.

@KorvinStarmast Okay so let's assume that the engineers somehow negotiate that hot air supply through the wing. Now how would it reduce wingtip vortices? am I missing something?

@AdityaSharma Yes, you are missing something. I was directing a comment to RalphJ. And no, handwaving that bit isn't a valid assumption.

@KorvinStarmast I'm sorry, I meant that everyone here including you is talking about the logistics of having hot air supplied to the wingtips, but I don't get the aerodynamics of it. How does hot air reduce wingtip vortices to begin with? Nobody seems to be questioning that, so I got confused and asked you. Good day!

Just a short "feasibility study", off the top of my head:

To resist such high temperatures, the structure should be made of some exotic titanium or stainless steel alloy which would add cost and possibly weight i.e. drag.
The exhaust pipe spraying hot gases would add complexity and weight i.e. drag.
Removing the low pressure (suction) from the top of the wing would simply kill the lift there since the biggest part of the lift just derives from the underpressure over the wing.
If the exhaust still posseses high pressure, this can be converted in additional speed and thrust simply prolonging the exhaust (convergent) pipe.
Tip vortex forms... well, at the tip of the wing; if they form away from the tip they wouldn't be tip vortex anymore :)

Anyway, tip vorteces are not that evil phenomenon as normally presented by the (marketing departments of the) airplane's manufacturers. Without entering into much detail (there are already many good explanations here about it), tip vorteces are just one piece of the big puzzle about a finite-span wing generating lift; if a wing is generating lift then it is also generating vorteces, due to Newton's second and third laws (aka momentum conservation). Indeed, vorteces are released downstream from the entire trailing edge of the wing, not just from the tips. The following picture (source) schematically represents this:

In particular, the strength of each of these vorteces is proportional to the spanwise variation of lift. For a conventional jetliner trying to reduce drag, lift is distributed from tip to tip with a shape as elliptical as possible in order to minimise that part of the drag termed induced drag (aka drag due to lift). Which such elliptical spanwise lift distribution, the biggest variation of lift happen towards the tips and that's why the biggest vorteces form at the tips. Anyway this is nothing bad, on the contrary it actually shows that the lift is distributed in such a way as to minimise drag.

I still don't get how hot air will reduce the wingtip vortices . Would it somehow remove the low pressure over the wing? (if that's what is to be achieved, you might as well remove the wing from there)

@AdityaSharma: well, I only did a short feasibility study; if it really happens, that I don't know ;)

Alright, good day!

To get enough flow to make a difference somehow, you might as well just point it aft with a nozzle to offset the drag with added thrust instead.  Like, say, the engine's own tailpipe.

@JohnK, I was thinking the exact same thing!


text:
Was there any attempt on making a turbofan for mach 1-2 with a "decelerator" air intake?
What made me ask this question is the SR-71 hybrid engines and the efficiency difference between certain engines.
Here is an article from wikipedia showing the efficiency of the engines accordingly to mach speed:

The air intake of the SR-71 engines had the job of slowing down the air flow so it wouldn't damage the engine blades, and it also used some of this air stream back to the combustion chamber; making it a hybrid between turbine and Ramjet engines.

So I was wondering if such thing could benefit Turbofans in higher mach speeds, not as extreme as the SR-71 turbines, but at some extent. It would be interesting to see a Turbofan at such higher speeds maintaining the the same level of efficiency.

Of course, the question is: "has this being tested/done before?", not if it is possible.
And if the answer is "no", then this idea is probably not possible/viable/practical.

The air intake of the SR-71 engines had the job of slowing down the air flow so it wouldn't damage the engine blades.

The main job of an inlet is to reduced the speed of the incoming air to some Mach 0.4, whatever high the speed of flight is. But not because that would damage the blades, rather because the range from Mach 0 to 0.4 is the limit within which the compressor's blades (and especially their tips) work in a proper way.
If the speed of flight is higher (Mach 0.4 to 1), then the inlet simply works as a divergent inlet (diffuser): its section gets bigger and bigger, the pressure gets higher and higher and the speed gets lower and lower (the compressor thanks Bernoulli).
And if the speed of flight is even higher (i.e. supersonic), in front of the inlet one or more shockwaves form, across which the speed reduces to high subsonic and afterward a divergent portion of the inlet finishes the work reducing again the airspeed till Mach 0.4.
This is how inlets of a modern jetliner, the Concorde or the Blackbird work(ed) to supply their compressors with a relative slow and stable airflow.
Note that the airspeed seen from a blade is the sum between the aircraft's speed and its rotating speed and it might indeed becomes supersonic: this can happen for example on the blade's tips of the fan of a modern turbofan.
Supersonic speeds on blades is normally avoided due to the decrease in the aerodynamics performance: lift decreases and drag is higher as well as noise, as visible for example in the following plots from this NASA report:

And even if, just for fun, we would design a compressor with supersonic blades, after it  the combustion chamber comes: combustion is an extremely slow process, so slow that burning velocity is normally measured in cm/s (inches/s). So after our supersonic compressor we would have to slow down anyway the airflow to low subsonic speeds.
Or we switch to a (sc)ramjet, but that's another kind of machine.


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There is any kind of advantage and disadvantages in making the blades of a helicopter off-center?
So, english isn't my first language, so I probably couldn't find any kind of information about it due to not being able to define what I'm asking.
Anyway, helicopter blades are normally like this:

An "off-center" helicopter blade would be like this:

I'm quite ignorant about aerodynamics and I couldn't find anything about it, so I hope someone here knows if a design like that would have any kind of advantages whatsoever.

Given the physics involved, they will try to straighten themselves out, putting a lot of stress at the 90 degree joint, causing that joint to either fail or have to be very over-designed.

Plus it’d make controlling collective and cyclic pitch much more difficult.

Just curious, but what do you think the advantages might be?

There is a theoretical advantage of an L shaped rotor blade insofar as it behaves like a swept wing, and so it's able to operate at a higher mach number where its velocity is highest, at the tip.
The problem is the blade as pictured would need to be massively heavy to be able to resist centrifugal forces trying to straighten it, and the torsional forces trying to make it twist it in ways you don't want.
So, since we only need that sweep angle where the velocity is highest, at the tip, we simply sweep the tip of the blade instead of from the root, like BERP tip.
So what you are showing is actually being done, but just with the dogleg in the blade swapped out to the other end since that is the zone that benefits from it, although even with that tip, as pictured below, you will notice the entire planform is shifted forward to place the mean center of lift of the swept section closer to that of the straight section, to minimize the torsional and bending side effects of the offset.


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What make and model are these two small propeller planes?
What are the makes and models of the two planes in the picture? The registration numbers are ZS-JNJ and ZS-LCY, the year was 1981 and the place was Botswana.

They're a couple of Cessna 210s. I can't tell from just looking whether they're Turbo 210s or normally aspirated. You can tell they're not 172s or 182s because they lack wing struts. Cessna 177s also do not have wing struts, but their front windshields are considerably different.

I (partially) own a 177 and I will back you up saying that these are definitely not 177's.

Yes, definitely 210. And not the pressurized P210, as you can tell from the windows. At first glance almost looked like 205s, or 206s, but the cantilever wing, retract gear, and window configuration give it away

These may be Cessna 210 because of their looks and the visible sections. It is a turbo propeller aircraft with one engine.

While I believe there was/is a turbine powered conversion available, the 210s pictured probably had 285 hp Continental IO-520 engines. At least that's what the 210s had that I spent several hundred hours in.

Thank you,  I finally found a registration listing that shows ZS-LCY as a Cessna 210. ZS-JNJ is shown as a Cessna 310, but that has to be a re-use of the number. Thanks again for the clues.

Yes, C210s, but not necessarily equipped with turbochargers.


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Is induced drag not caused by tip vortices?
According to most online resources, including NASA, induced drag is caused by wingtip vortices which induce downwash. However, in StackExchange I have found a few answers (Peter's and Jan's) that actually drag caused by vortices is a myth and induced drag is a simple inherent reaction force (vortices and induced drag are both consequences of downwash, which itself is caused by lift creation).
Although this theory looks quite intuitive, I am still confused. Could anybody give me some links to more extensive research of this issue or explain which one theory is scientifically correct and why?

Welcome to aviation.se!

The danger in asking which one theory is scientifically correct is that questions about induced drag / lift have the Blind men and the elephant problem. Just remember that any simple answer to a complex question will be at least a little bit wrong.

Well, I understand that there is no simple explanation of induced drag, but the two theories contradict each other (at least it looks like). There has to be a lot of scientific research in this field, but I cannot find any extensive explanation about why induced drag is NOT caused by vortices.

Oh, but they are all* correct!
First, we have to say that we are not talking about cause here. There is a flow around the wing and it has certain properties and neither is cause of the other; they are both effects of the flow. What we are talking about here is explanations why every flow that produces lift also produces induced drag.
And as far as explanations go, there can be more than one, because the physical theories have a lot of redundancy in them. You can usually arrive at the result by multiple ways.
You can use the orientations of the normals (the Peter's answer), you can use laws of conservation of energy and momentum (my answer) or you can use the potential flow theory and that explanation will indeed involve the wake vortices. Because the potential flow theory is a mathematical apparatus for analysing fluid flow as a combination of vortices and from that angle of view, the energy is indeed carried away by the wake vortices.
However, the wake vortex explanation has one problem: it is often associated with different misconception, the one that the vortices are caused by transverse flow around the tip and the resulting idea that it can be eliminated by doing something with the tips specifically.
But that's not what the potential flow theory says. According to the potential flow theory, there is a vortex originating from every point along the span and they combine to form the two tip vortices. So like the other explanations, it also clearly says the induced drag is related to any generation of lift. And if you understand that, it is correct. I still consider it unnecessarily more complicated.

* All the three you listed that is. The ones that claim the wingtip vortex is caused by transverse flow around the tip are not, because it isn't. The wingtip vortex is just the vortex line of the overall wake (and should preferably be called wake vortex).

Now explain it for helicopters. 8^D

Aerodynamics is so simple when it's introduced this way.

@KorvinStarmast, the arguments apply to rotorcraft almost directly. The only difference is that the flow is due to rotation of the rotor instead of forward motion. But the lift generation works the same.

@mins, the detail are really, really complicated. But physics has many large scale laws (mostly based on conservation of something) that allow you to skip all the details and still get useful insight to the overall situation.

Thank you, Jan. It is getting more clear to me. But does it mean that winglets do not have any effect except for changing the aspect ratio a little bit?

@RokasZilinskas, winglets have a little bit of effect, but indeed not very big and there are better ways to reduce drag, so winglets are only used in upgrades to old designs, because they are easy to retrofit, or if the wing span is limited by other concerns. For more details, see this analysis by Peter Kämpf.


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How to record flights and cockpit radio com using an Android phone?
I am a student pilot, and would like to record my training flights (in a C152) to look back on. I do not have a Go-Pro, and do not want to fork over the money for one. I do however, own a Samsung Galaxy S5, that has a pretty good camera.
I was wondering if anyone had some ideas as to how I could record my flights using my phone? I have bought a mount, a Y-cable connection for the 1/4"" phone jack, an TRRS cable and a Headphone/Microphone splitter cable for recording the headset audio.
I have, however, noticed that the avionics interferes with the recording and causes the phone to stop recording. Does anyone happen to know why this is?
In addition, this question is not limited to my specific cables/connectors.
Please let me know if anyone has found a working solution with their Android phone in the past. A similar question was answered for IPhone users. Does anyone have experience with Android phones?

Whenever I turn on the avionics, the recording may stop and go to the "OK Google" feature but that is due to repeated talking in the microphone. I believe the avionics stops the recording as I have seen both audio recording and video recording apps get stopped during mid flight. Does anyone know why?

I cleaned up the links for you, but you should be able to edit your own question regardless of how much reputation you have.

OK thank you - stack exchange stopped me, it said I needed 10 rep points prior to posting. Perhaps afterwards I can...not sure.

When I used a camera that didn't have an audio input, I had a cheap handheld audio recorder from Walmart that I hooked into my headset jack. I later synced up the audio by giving a visual cue and saying something ("3..2..1..Mark" and using fingers). I then synchronized the audio in Windows Movie Maker. After that I decided to "fork over the money" for the Garmin VIRB.

@JonathanWalters new users are limited to 2 links per post, IIRC

@Ron Beyer, thanks for the suggestion. I was trying to avoid having to sync between two files. If it comes to it, I'll go your route.

Android by default has Google Now installed. Voice search listens for the words "Ok Google" when you speak, but phones also detect a long press in a headphone button (on headphones with controls). It can also be fooled into activating by odd signals like the pop when you turn on the avionics master switch or loud droning noises like motorcycle and airplane engines. [credit: @GdD 02/03/2017]
In this case, it was the static and popping noises made by transmissions of the radio com that would interrupt recordings as it masked itself as the long press on the headphone jack. This would stop a recording and start the voice search app.
A quick fix is to disable Google Now. This disables any Google features including "headphone long press" detection from the headphone jack. No other app uses the long press to activate, effectively eliminating the problem from ever interrupting your recording during flight. No need to download an app just disable Google Now in Settings > Applications > Google Now > Disable.
Once this is resolved, I noticed that my video won't record past 4 GB. This is because of the FAT32 file system limit of my SD card (default for Android SD cards). An app can fix this and automatically restart a new recording when the file system size limit has reached. There is no getting around this limitation of the file system other than restarting a new file.

On Marshmallow, if you make the SD card "adoptable storage", it'll use ext4 instead of FAT32, and you can have files larger than 4 GB, at the cost of not being able to take the SD card out of the phone and use it elsewhere. But at least on the S4, Samsung has put a limit in the camera app to stop you tripping over the 4 GB limit, regardless of what filesystem is on the SD card. See How to record More than 4GB Video on exFAT? on Android.SE.

@DanHulme: ext4 works on all (newer versions of) Linux distros, not just Android.

Android by default has Google Now enables, which include voice search. Voice search listens for the words "Ok Google" when you speak, but phones also detect a long press in a headphone button (on headphones with controls). It can also be fooled into activating by odd signals like the pop when you turn on the avionics master switch or loud droning noises like motorcycle and airplane engines.
First thing I would try is to turn off the voice detection, see this page for instructions. Now google voice should stop popping up if the phone is being fooled by noise.
However, that may not be the end of it if the phone is detecting static or other signals as a long headphone control press, in which case it may still interrupt your setup. In that case you will need to use an app to disable the headphone controls, the one I've heard of is SoundAbout, or your phone's distribution may have settings to let you do it.

@GgD I think you are onto something here. I've tried this out with a headset and the phone did exactly what it does on the airplane. I think we may have just found out the root cause, hopefully that was it. I will try and report back. Thank you!

Good luck on your training @nndhawan

@GgD please accept my edit so that all can benefit from the final solution.

That's not what edits are for @nndhawan, I suggest you put it in a comment or answer your own question.

Ok @GgD. Please don't mind me if I take some parts of your answer. The insights you provided were accurate and helped me discover the solution. Thank you.

I don't mind you using part of my answer @nndhawan, but it is considered good form to cite it in your post. it's also good form to award answers to the post that helped you, rather than your own, especially when the other post gave you the answers you needed.

@GgD I want the answer that is the solution of the question proposed to be awarded the answer because that is what the question is in reference to. If I make your answer the top answer on the basis that you pointed me in the right direction doesn't suffice as an answer but as an intermediary step. We want the final solution to be considered the top. I tried to edit your answer but you told me not to do so. That is not my fault. In all honesty the answer was not complete so you don't get best answer, but you will get credit in my post.


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Why don't contra-rotating propeller configurations with more than two propellers exist?
I wasn't able to find a single example of more than two propellers stacked axially in a contra-rotating configuration. There is a huge number of contra-rotating propeller setups, but it is always a combination of two propellers on the same shaft, not four or six.
Why isn't a piston engine used to run more than two standard propellers on a single shaft? Hasn't it been tried (or tested at least) in 100+ years of aviation history?
(This question isn't about turbofans/turbojets, but I also couldn't find anything for ducted fans: two props on a single axis is a maximum.)
Blade-solidity is explained here:
Contra-rotating props explained here:
An increase in a number of blades does seem to reduce the efficiency, whereas an increase in propellers on an axis seems to increase the efficiency.

shaft design. It's much more complicated the more moving shafts you need to embed within each other

What would be the advantage of this? There are only two directions each propeller can spin, and why would one want to have several propellers on the same axis spinning in the same direction, rather than just one propeller with more blades?

I've added two references into my original question. Increasing the blade solidity ratio is the worst technique one could employ for packing more thrust into the same space, on the contrary to Contra-roration which seems to give some benefit.

I would guess it is diminishing returns. The first CR prop gets you 6-16 improvement, but I doubt the next one gets you that much and the mechanical complexity has to get totally crazy with three or more shafts rotating inside each other.  So the guess is that it doesn't help very much and costs a lot.

I would think that in a rotary wing aircraft especially, a four rotor stack would not only be obscenely complex, but in terms of the stresses involved on the entire structure it would seem likely that it would have to be a low maneuverability vehicle.  Else you would risk fracturing the stack and becoming a rock.

No, stacking more propellers will reduce efficiency. The case with two is an exception because the second propeller benefits from operating in the slipstream of the first. In total, both propellers accelerate the air backwards without adding swirl losses.
However, propeller efficiency goes down with increasing flow speed ahead of the propeller. This means that every other propeller operating in the accelerated flow of a preceding propeller will have lower efficiency. Generally, it will be better to spread the propellers out sideways than to put them in line.
Note that contra-rotating props are the exclusive realm of very powerful engines with very high disk loading. The most recent designs prefer to use more blades and avoid the heavy gearbox, but the efficiency of such props is poorer than that of propellers with lower disk loading and fewer blades. In a way, the efficiency of the eight-bladed propellers below is similar to that of four two-bladed propellers in sequence, but it is much easier to build with all blades and their pitch mechanism in one hub.

T-56 engines with Hamilton-Sundstrand NP2000 propellers on an LC-130. Photograph by Robyn Waserman, National Science Foundation (Date Taken: November 21, 2008)

In general, isn't it preferable to have a blown wing, and the more of the wing you can get, the better? The added complexity (and weight) penalties that you get from contrarotating props seem to outweigh the gains, except perhaps in a narrow range of scenarios.

@egid: The blown wing only helps during take-off and landing, but is a source of added drag in the time between the two. The higher flow speed and turbulence in the slipstream creates more viscous drag.

Ah, interesting. Makes sense.

Weight of gear train and power required to drive more than a pair of contra-rotating props makes two the max.  It was thought of and disposed of a long time ago.

Well, if you locate that additional propeller(s) in another place on the airplane then you need an additional gearbox+shaft+possibly another engine. I don't think this weight is going to be less than just a slightly longer and stronger shaft.

@sophit From what I understand, simply extending an existing shaft is not enough to accommodate a third propeller. Here is a diagram of a contra-rotating propeller assembly; shaft-1 cannot be extended forward because shaft-2 is in its way. If we were to extend shaft-2 rearward to drive a third propeller behind the other two, then it would be impossible to power the shaft-1. It seems that the only way to accommodate a third propeller is to have a third shaft.

@AdityaSharma yes sure, one shaft for each propeller just like in a multistage turbojet there's one shaft for each stage. Or, you attach two propellers on the first shaft and two propellers on the second one. Just speaking about weight here, aerodynamic or aeroelastic phenoma have been already dealt with in another answer.

One shaft per propeller seems like the only option here... Regarding your latter suggestion, two propellers per shaft are probably going to function like a single propeller with just more blades, so I wouldn't consider it as four discrete props (but I might be wrong on that one).


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How does hovering in ground effect (IGE) reduce drag?
I understand that when IGE the induced velocity and thus the induced angle is reduced, rotating the total aerodynamic force (TAF) closer to the vertical and hence increasing the vertical component of lift (for the same power).
What I don't understand is why drag is reduced (some texts refer to a reduction in induced drag, others to rotor drag). Diagrammatically, the horizontal component of drag (parallel to the tip path plane and that which opposes rotor RPM) actually appears to increase.
I've concocted this diagram which may itself be incorrect and the source of confusion, or maybe its just semantics?

@mins Actually (for the same power) the reduced induced wind velocity serves to increase (not decrease) the AoA... that's where the extra (vertical component of) lift comes from. Like lift, drag has only changed in its applied direction, but not in quantity. That's what I'm trying to resolve here, why is the now more horizontal drag not disadvantageous? I would particularly appreciate answers that relate to my pretty picture, 'cos that's how my brain works. Thks.

@mins Agreed. So is THAT why drag is said to reduce? The reduced pitch angle means less everything: drag and lift! The reduced lift is only possible because the lift vector is more vertical, therefore proportionally more of the lift vector is contributing to oppose weight. I find that saying less power is required in the hover because of a reduction in drag to be a little misleading. There is less drag because you lowered the collective! I think it is more accurate to say that there is an increase in lift due to the (un-commanded) increase in the AoA, which leads to less collective, drag etc.

I find that saying less power is required in the hover because of a reduction in drag to be a little misleading But it's true. Lift in a helicopter in powered flight does not happen in a vacuum.  It happens in the context of the engines providing the power needed to create the rotary wings' velocity that then creates the lift that keeps you airborne, in or out of ground effect.

The inflow becomes less vertical and also reduces in velocity since the pressure under the disk increases.  The rotor airspeed is therefore reduced and induced drag reduces in proportion to the square of the reduced airspeed.
The tip vortices (recirculation) are also reduced meaning that the outer portion of the blade, where most of the lift is generated, becomes more efficient.
Since the angle of attack has reduced, because the inflow is more horizontal and the lift vector is now more vertical, you get more lift for the same pitch angle and therefore require less power.
Less power into the rotor = reduced induced drag.
This image, from Copters.com, may help.

To add some clarifications.
Why does the rotor effective airspeed reduce?
Imagine a helicopter in a zero wind, ideal conditions out of ground effect hover.  The rotor, to produce lift, accelerates air downwards. Since there is no wind and the helicopter is not moving, the air is drawn into the disc from around it and mainly from above it.  Ignoring side effects, the air is drawn down equally all around the disc.  The airflow drawn into the disc has a horizontal component and a vertical component.  The vertical component is large as there is little to stop the disc from pulling air down and accelerating it downwards.  The speed of the rotor plus the horizontal component of the inflow co-incident with the average chord of the blade can be thought of as the "relative airflow".
Now imagine the same except hovering in ground effect.  The cushion (increased pressure) below the disc reduces the inflow (imagine a fan partially blocked, the fan draws in less air) and thus the horizontal component of the inflow is reduced.  Therefore, the relative wind airspeed is reduced.  As you can see from the diagram solving the triangle of vectors, the angle of attack is also reduced.  The collective could actually be reduced to accelerate the same amount of air.  This power reduction results in a decrease in the induced drag.
Why does the same pitch angle produce more lift when hovering in ground effect?
The biggest "drain" of power when hovering, and the main reason why hovering uses more power than any other flight phase, is the vortices around the rotor tips and the outer third of the rotor where most of the lift is produced.
When hovering out of ground effect, there is nothing to stop the accelerating column of air under the disc from "escaping" and being drawn back up into the in-flowing air leading to re circulation and significant vortices.  When hovering in ground effect, the air is forced to move out laterally from under the helicopter which effectively increases the radius of the vortices.  There is a significant reduction in the turbulence around the tips and the outer third with a corresponding large increase in efficiency.  Therefore, more lift is generated in ground effect than out of ground effect for the same angle of attack.
This can be seen practically by hovering over a hard surface and measuring power required then comparing this when hovering over long grass or worse, hovering over a dense standing crop.  Since the long grass or crops reduce the ability of the air to escape further out laterally, the vortex radii are not increased as much as when hovering out of ground effect and more power is required.
It is possible, when there is little power in reserve, to be unable to take off vertically from long grass but be able to do so from the adjacent hard standing.
I think this video helps to visualise the vortices.

Can you please elaborate how the reduction in vertical inflow results in The rotor airspeed is therefore reduced? I would have thought that the reduction in vertical inflow results in (if not compensated for by the pilot by lowering the collective) a higher AoA which results in both more lift and more drag? Also, you state more lift for the same pitch angle - but this also means (by definition) that the AoA has increased not decreased.

@jumblie I've clarified with an edit.

Inadvertently, the question posed has less to do with ground effect and more to do with the total aerodynamic force (TAF) and vector diagrams.
The total aerodynamic force (TAF) can be resolved into different components dependent on the chosen frame of reference. The frames of reference most useful in studying aerodynamics are the;

resultant airflow (RAF); the TAF resolves to lift (perpendicular to RAF and drag) and drag (perpendicular to lift, parallel and opposite to RAF)
plane of rotation; the TAF resolves to rotor thrust (perpendicular to the plane of rotation, blade span and rotor drag) and rotor drag (parallel to plane of rotation, opposing rotation; perpendicular to rotor thrust)

The original diagram in the question resolves the TAF relative to the RAF and fails to differentiate between blade drag and rotor drag. The benefits associated with hovering IGE come from the reduction in rotor drag (less power required).
To understand this by studying the vectors, the TAF should be resolved with reference to the plane of rotation (not the RAF). In this way, it is quite clear the rotor drag reduces when hovering IGE due to the more vertical total reaction which results from the reduction of the induced flow.

I would have made the diagram for constant thrust instead of constant TAF. The effect is then more pronounced and more accurately answers the question, which is about hovering (requiring constant thrust). Still, +1 for a succinct answer.

Induced drag is caused by bending the air downwards to create lift: the higher the deflection angle, the more induced drag. From the wiki for induced drag:

When in ground effect, the air cannot deflect very much, but bounces back off the ground. It cannot penetrate the ground, which acts like a mirror, creating a higher pressure underneath. Part of the lift is now created by higher static pressure under the rotor blades.
So Out of Ground Effect OGE, all lift is created by deflection of the air mass, and deflecting at a lower angle results in less lift. IGE, the ground mirror effect pairs lower angle deflection (less induced drag) with a higher static pressure underneath the rotor (same amount of lift).

An IGE hover, the downwash hits the ground, creating a cushion of air. This cushion of air reduces the amount of power required to remain in the hover, reducing cyclic, and therefore reducing drag.